UNCLASSIFIED

275",463Al" ,,,

ARMED SERVICES TECHNICAL INI'MMON AGENCYARLINGIDN HALL STATIONARLINGTON 12, VIRGINIA

UNCLASSIFIED

Di• NOTICE

0-0

THIS DOCUMENT IS BEST

QUALITY AVAILABLE. THE COPY

FURNISHED TO DTIC CONTAINED

A SIGNIFICANT NUMBER OF

PAGES WHICH DO NOT

REPRODUCE LEGIBLY.

NOTICE: ihaen governwt or other dravings, speci-fications or other data are used for any purposeother than in connection with a definitely relatedgoverment procurame. operation, the U. S.Government thereby incurs no responsibility, nor anyobliation vhatsoever; and the fact that the Govern-ment my have forimlated, furnished, or in any vaysupplied the said drawings, specifications, or otherdata is not to be regarded by implication or other-wise as in any manner licensing the holder or anyother person or corporation, or conveying any rightsor permission to manufacture, use or sell anypatented invention that may in any vay be relatedthereto.

( '

THE JOHNS HOPKINS UNIVERSITY1I L APPLIED PHYSICS LASIRATGRY CM-1010

121 k.gie Aen. Silver Sp*n. UhrgiuW

Operating under Contract NOrd 7386 with the 4 -{ Burpou of Naval Weapons, Department of the Navy Copy 4IO. JL.)

SECONDARY GAS INJECTIONIN A CONICAL ROCKET NOZZLE

I. EFFECT OF ,R!FICE DIAMETERAND MOLECULAR WEIGHT OF iNJECTANT

byR. E. Walker, A. R. Stone, and M. Shandor

oaSe4 to IS 1 bY 'VALBureau Of I Ji' V ,

February 1962

CM.1010

Februry 1962

Secondary Gas Injection in aConical Rocket Nozzle

I. Effect of Orifice Diameterand Molecular Weight of Injectant

byR. E. Walker, A. R. Stone, and M. Shandor

fHE JOHNS '4OPKI S uNlVfmS TY

APPLIED PHYSICS LABORATORY46 I B-A A llVI II rN G M A IV AN

w

TMe Jshri ""eking UANyW6tytAPIM1 PU"IWo UATBy

ABSTRACT

Data are presernted on interference forces result-

ing when a gas at arbient temperature is laterally in-

jected thi'ough a single circular orifice in the conical

portion of a rocket nozzle into hot supersonic propellent

gases. Variables cxamined are (a) injectant orifice sizeand associated pressure ratio change and (b) effect of

injectant molecular weight and specific heat ratio.

Other parameter; remain essentially constant during these

tests. It is shown that simple theoretical arguments

can predict relative 3ffects of intrinsic injectant

properties, but that pressure ratio effects are not ade-

quately described. It is also shown that effectivenessof secondary inj':ction depends on injectant orifice size,

which has not been treated in Any theoretical models.

- ii -

*FftS5S Pum Ia. T.The Joims WOWie Ukws~it

TABLE OF CONTENTS

List of Illustrations iv

List of Tables v

List of Symbols vi

I. SUMMARY ND) CONCLUSIONS 1

II. BACKGROCID . . 3

I11. DESCRIPTION OF EXPERIMENTAL APPARATUSAND TESTING PROCEDURE .. 5

IV. EX)?-.,.SMENTAL RESULTS . 9

V. THEORETICAL FOUNDATION . 13

vI. DISCUSSION OF THE DATA . . 18

References 41

Acknowledgements . 44

- iii -

[1

LIST OF ILLUSTRATIONS

Figure Page

1 Research Rocket Motor Used in SecondaryGas Injection Experiments . 34

2 General Setup of Apparatu3 for SecondaryGas Injection Experimerts a.

3 Effect of Orifice Area on Secondary GasInjection (CO2 Injectant at 700F) 364 Correlation of Secondary Injection Data

for Varioua Injectant Gases(d. = 0.0625 in., Toj = 70F) • 37

5 Secondary Gas Injection Model 386 Approximate Pressure Rise Due to Induced

Shock Wave (Subsonic CO2 Injection Data) 397 Correlation of Sonic Secondary InjectionData [CO2 Injectant at 700F; All Sub-

sonic Injection Data (PoJ/Pl 4.0)Are Flagged] .. 40

- iv -

r -)Oh -Oif uk

LIST OF TABLES

Table Page

I Properties of Research Rocket Motor,Nozzle, and Injectants Used inSecondary Gas Injection Experiments 22

11 Secondary Gas Injection Data 23

III Some Properties of the Injectants Use. in

Secondary Gas Injection Studies .33

i A4 ~h p~.l uIat.O d~OftJom w f eakwmaf

LIST OF SYMBOLS

A = Jet orifice area

At = Nozzle throat area

A, = Nozzle area at point of injection

CD = Discharge coefficient of orifice assuming sonicflow

d = Orifice diameter

FN = Side force due to secondary injection

F = Axial thrust of motor

I s .FN/Wj, Effective specific impulse of injectant

1 8 Specific impulse of vacuum-exhausted sonicjet uf injectant, Eq. (7)

= ach number of nozzle flow at point of injection

Mach number of jet gases after expanding tofreestream pressure

= Molecular weight of propellent gases

ij = Molecular weight of injectant

S1(Mj) = Mass flow function defined by Eq. (5)

Pl = Static pressure of nozzle flow at point ofinjection

P2 = Static pressure behind induced oblique shockwave

P = Stagnation pressure of propellent gases

Poi = Stagnation pressure of injectant

P = Static pi-essure of injectant at orifice

vi

The M- OuiWkZlU:y

APPIS PbSe.g Ufaewy

T 0 Stagnation temperature of propellant~oSTo= Stagnation temperature of injectant

j = Mass flow rate of injectant

V =Mass flow rate of propellant

Y= Specific heat ratio of propellant

Y = Specific heat ratio of injectant

= Conical nozzle half angle, 150

- vii -

The Joiwd U0inm UMaWhy~~ArPL190 POlrIM LAIWATS

SECONDARY GAS INJECTION i.. A CONICAL ROCKET NOZZLE

I. Effect of Orifice Diameter and Molecular

Weighc of Injectant1

I. SUMMARY A.) CONCLUSI3NS

This report is the first in a series dealing with an

experimental study of thrust vector cintrol using gaseous

secondary injection. For this study. t.e main propellant

was a hot gas (catalytically dfcoapose'd H202); a variety of

gases (CO2, N2 , Ar, 0.8 He + 0.2 Ar, He, and H2 ) at ambient

temperature was used as the injcta.it. A conical convergent-

conical divergent exhaust nozzle was used, with injection

normal to the nozzle axis at a fixed ooint in the divergent

portion of the nozzle.

A variety of circular orifi-,-; diimeirs (0.180, 0.125,

0.089, 0.0625, and 0.04 inches) was examined. The side force

developed by secondary injection was .easured directly with

a force transducer; the data are reported as specific im-

pulse ratio or "amplification factor" obtained by dividing

the measured effective specific impulse of the injectant by

the specific impulse of the injectant for sonic flow into

a vacuum.

As thu orifice diameter was varied (ith CO2 in-

jectant). low-pressure injection was critical)y examined.

The results showcd that for a particular ori!'ice size the

IThis work was sponsored by the Special Project-, Office,Bureau of Naval Weapons.

The. Jdww -ah UftWfAPMINO FRI,6 L A TOuY

amplification factor has a maximum at or near the transition

from sonic to subsonic Anjection. Performance does not in-

crease indefinitely for decreasing pressure ratio across

the orifice, as might be construed from simple linear super-

sonic flow theor. Significant effects of orifice size on

the specific impulse ratio were observed: For a fixed

pressure ratio across the orifice, performance increases with

decreasing orifice size.

Subsonic injection data aided in estimatii.g the

strength of the shock wave induced in the supersonic flow.

These data have shown indirectly that shock wave strength

increases to a limiting value close to that required for

shock wave-turbulent boundary layer separation.

Performance of several inert gas injectants with

differing molecular weights and specific heat ratios corre-

lated well with a parameter suggested by linear supersonic

flow theory. One potentially reactive injectant (H2 ) was

used; its data correlated well with the inert gas data and

suggest that essentially no reaction occurred in the nozzle

between the injectant and propellant.

Additional experiments designed to measure the effects

of other parameters (such as injectant temperature, motor

temperature and pressure, and injection and nozzle geometry)

are desirable to establish appropriate theoretical avenues.

Some of these experiments are presently in progress at this

Laboratory.

-2-

the Johm H"p&Itn UnivwillyAPPLIES POYBISS LA SOATINY

akIu w Ing. *atyIe~

II. BACKGROUND

Recent advances in solid rocket propellant technology

resulting in higher flame temperatures and multiphase flow

have increased the desirability of thrust vector control

methods that do not require exposing moving material parts

to propellant exhaust products. For this reason, other

methods of deflecting the supersonic nozzle flow are being

examined. Secondary injection Js one method that has received

considerahle attention. This technique utilizes the forces

developed on the wall of the divergent portion of the rocket

exhaust nozzle by lateral injection of a fluid (gas or liquid)

into the supersonic propellent gases. In addition to the

usual jet reaction, local high pressures associated with an

induced shock wave "amplify" the jet reaction. The first

exrerimentm on secondary injection were reported by Hausmann

(1)2 and demonstrated that the shock-induced reaction asso-

ciated with an air jet directed into supersonic air (both

gases at ambient temperature) could for certain conditions

be as large as the jet reaction. The nature of this shock-

induced reaction and how It depends upon the mainstream and

i ,joctant properties has since been the subject of considerable

study, mostly experimental. Several experiments have been

reported for Jet-interference phenomena on simple aerodynamic

surfaces (2-6), on the external surfaces of simple missile

configurations (7-9), and on the internal surfaces of rocket

nozzles (secondary i:.jection) (10-16). Most of these reports

2 Numbers in parentheses indicate references at end of' paper.References are listed on pages 41, 12, and 43.

- -

The JAM H@eOkias UiovwsityAPPOIII Phyll LA"RATINV, 4hvr IpAa M~fyrun

deal with ambient temperature air-air interaction. Becauseof the cohiplexity of the problem, analytical descriptionshave been quite limited. For gas injection, two qualitatively-correct basic descriptions have been useful (8, 17-19), butrefinements are desirable.

This report presents the results of experiments per-formed at tis Laboratory on secondary gas injection intohot supersonic propellant flow in a small rocket motor.Data on the effects of injectant gas properties and injectantorifice size have been obtained; motor operating conditions,nozzle geometry, and injectant location have been kept fixed.Future experiments involving temperature effects, nozzlegeometry, point of injection, and mainstream properties areplanned.

-4-

the JohnHopin ow&IUiiltyAPPLIlS PUY$#" UNATORV

III. DESCRIPTION OF EXPE!IIMENTAL APPARATUSAND TESTING PROCEDURE

Agparatus

The data presented in this report were obtained with

a Tmall research rocket motor and nozzle, sketched in Fig. 1.

The working fluid wag provided by catalytic decomposition of

90 per cent hydrogen peroxide liquid3 at a nominal motor

chamber pressure of 400 ps&. The products of decomposition

were 29.2 per cent mole fraction of oxygen and 70.8 per cent

mole fraction of water vapor, with a specific heat ratio of

1.266 (20). Average propellant exhaust temperature, measuredwith an uncalibrated Xron-constantan thermocouple, was 1845°R

(Rankine), with a maximam spread of 1830-1865°R. This tempera-

ture is slightly higher than the theoretual adiabatic de-

composition temperature, 182503 (20).

gome motor, nozzle, and injectant properties are listed

in Table 1. Attempts to measure the liquid propellant flow

rate * were unsuccessful. It was therefore necessary to

rely upon motor chamber pressure P and temperature T (=18450R),

geometrical area of the nozzle throat At(=0.196 in ), and

isentropic flow relations to compute propellant flow rate,

theoretically, W 0.00280 P0 where f is in lb/sec and P0 is

in psia.

The average thrust coefficient C. for motor chamber

pressure of about 400 psia has been found experimentally to

be 1.42, which is somewhat below the theoretical value of 1.46

3Supplied by Becco Chemical Division, Food Machinery and

Chemical Corporation.

-5-

The Johns H&ins UnIvers~vAMPI|D P1"IlN LAMNATIN

56ySInk. #"I"n

(based upon conical isentropic nozzle flow and area ratio

for P = 400 psia, and an atmospheric pressure of 14.7 psia).

Axial thrust F for these experiments can be closely computedfrom F =-. CF A tP 0.278 Po, with F In lb and P in psia.

Injectant gases were obtained from standard compressed

gas containers.4 A maximum injectant pressure of about 500

psia was used for carbon dioxide and about 1000 psia for the

other gases; injectant gas temperature was ambient, nominally

7007. The injectant gas flow rate was meterd thiough a

standard ASME sharp-edged orifice tlowmeter calibrated with

CO2 by timed discharge into a calibrated volume. Molecular

weight and compressibility corrections were made in the usual

manner when other injectant gases were used. Evaluation of

the discharge coefficient of the injectant orifice in the

nozzle wall by using a combination of this metered flow rate

*j, measured jet total pressure Psiq total temperature Toj

geometric orifice area A (=TTd 32/4), and isentropic flow re-

lations provided a compatibility check for several measured

parameters.

Figkre 2 shows the general setup of thc apparatus.

The motor is mounted on the periphery of and in line with the

axis of a drum, the axle of which is mounted in antifriction

bearings that permit simultaneous rotation and axial motion.

The drum floats in water to reduce bearing load. Force trans-

ducers measure axial motor thrust and turning moment developed

by secondary injection or by motur trim misalignment. Pro-

pellant and injectant go to the nozzle through relatively

long rigid lines which by test were found to introduce fixed

4 Carbon dioxide for these experiments was supplied by PureCarbonic Company. All other gases wexe supplied by SouthernOxygen Company.

the Johns Hopki Un|ersity

APPLIIDPIsmI LA@*MATIeY

spring constants superimposed on the elastic constants of the

force transducers. Transducer calibrations are obtained after

or during each day of operation with the transducers in place.

Most pressure measurements were made with a variety ofelectrical pressure transducnrs which had been pbriodically

calibrated with bourdon element test gauges. The gauges hadin turn been calibrated against a standard dead weight tester.Where possible, all transducers were excited from a common

monitored supply voltage, and observed variations in excita-

tion voltage were included in the data analysis. Temperature

measurements were made with uncalibrated iron-constantanthermocouples. Appro::imecely attenuated transducer and thermo-

couple outputs were recorded on four 0-1 my. 10-inch Westronix

strip-chart recorders either continuously or through a dual6-point data sampler which permits more than one bit of in-

formation per recorder channel.

Testing Procedure

Because there had been some transients, the followingoperating sequence was adopted:

(1) One complete data sampling sequence (:12 sec),

without propellant or injectant flow, to establish transducer

and recorder zeros. No special effort was made to preadjust

transducer outputs to zero.

(2) Propellant-on--injectant-off sequence to deter-mine thrust misalignment (motor trim).

(3) Propellant-on--injectant-on to measure secondaryinjection effects.

-7o

IM JehA& foekina UAnwityAPPLIIE PNVSICS LABOIATORV

(4) Repeat of (2.) to determine trim change, if any.

(5) Repeat of (1) to determine transducer zero 3hift,

if any.

This procedure permits all bits of information to be

extrapolated and evaluated at a common time.

-8-

The isW. HeAk.P unnWa.yAoftUe3 PRY"" LAGONATORY

IV. EXPERIMENTAL RESULTS

Data were obtained on the separate effects of

(1) injectant orifice size, and

(2) the effect of injectant gas type.

Only injection normal to the nozzle axis was examined.

The motor chamber pressure was kept at the experimental maxi-

mum value of about 400 psia. The propellent gases had a stag-

nation temperature close to 18450R.

The injectant port, a single circular orifice, was

located at the point in the conical expansion nozzle where

the Mach number, M, was 2.34. This Mach number was deter-

mined both by the experimentally measured pressure ratio

PI/P0 = 0.0730 and the geometrical area ratio A,/At = 2.597.

The static pressure of the undisturbed supersonic flow at the

injection point was nominally 30 psia. The exit Mach number

of the nozzle M3 was computed from the geometrical area

ratio A3/A t to be 2.83. The exhaust gases were slightly

overexpanded at the nozzle exit, P3 = 12.8 psia. No attempt

was made to reduce P in order to examine secondary-injection-0

induced separation effects.

In the study of the effects of orifice size, carbon

dioxide at ambient temperature (nominally 7001) was selectd

as the injectant and the orifice diameter d was varied from

0.0625 to 0.180 inches. Changes in the orifice size were

accomplished by simply "drilling out" the just-tested orifice.

This practice gives a kind of thick, square-edged orifice

-9-

The J hmHi" UMfet.i , tyAPPUIB PHYA" &AMATORY

IF# SUFng. MwyISn

whose discharge coefficient can be expected to depend strongly

on Reynoids number and pressure ratio across the orifice. The

variation in this pressure ratio P oj/P brought about by

varying the jet mass flow was sufficient to give both subsonic

and sonic flow through the orifice (1.4 <Poj/PI 1 2). In retro-

spect, carbon dioxide was a poor choice for a working gas since,

at the pressures and temperatures involved, significant com-

pressibility elfeLtki were encouptered. (For example, at 500

psia and 70'r the compressibility factor for carbon dioxide

is 0.79 and represents a considerable and measurable deparcure

from ideal gas behavior). Compressibility effects were taken

into aceounl' when evaluating the orifice discharge 'oefficient

by using a lnearized treatment given by Eggers (Uxi for a

calorically perfect but thermally imperfect gas. Seixratit

exper.ents were performnd to establish the validity of using

this ].inearized analysis. No other compressibility effect

corrections were made to the data.

A detailed listing of the experimental data is provided

in Table I1. A summary plot of the data pertaining to effect

of orifice size is provided in Fig. 3, where the normalized

sp- ific impulse Is /I and sonic discharge coefficient VD

are plctted as a function of jet pressure ratio j/P I.

The effective upecific impulse I s is obtained by

dividing the force normal to the motor axis Pq by tLe 7:esured11

jet mass flow W.. I is the specific impulse of a Aonic jet

of the injcctant exhausting into a vacuum. I /I , therefore,represt n1- an amplification factor for secondary injection.

The sonic discharge coefficient CD is obtained by

dividing tue measureli jet mass flow by a theoretical value

based upon sonic isentropic flow, the geometrical area of the

orifice, and measured values of P03 and To. The breaking

- 10oJ

-10 -

14ih. M-0 1.IPA UgOIWtstyAPftSI6 PNHIGS LAWMATORV

away of CD from a constant value near unity is interpreted

to be a transition from sonic to subsonic injection. As seen

in Fig. 3, the knee of the CD curve occurs at a value of

Poj/P 1 larger than te critical pressure ratio for jet flow

without supersonic cr.ssilow interference. This results from

higher effective back pressures brought about by the induced

shock wave. Note also that the Is /I curve tends to peak

at or near this transiticn point and does not exhibit a mono-

tonically increasing behavior for a decreasing Poj/Pl, as

might be inferred from simple theory discussed later.

Finally, the itrong depende,:ee of secondary injection

effectiveness upon -rifice size shculd be recognized. The

variation in d examined here exceeds that studied by others

and the consequence of varying 4 has not been pointed out

before.

Several gases have been used to investigate the effects

of injectant molecular weightA and specific heat ratio N ;

the gases used are listed in Table III. All of these injec-

tants are inert with the exception of H2 which, in principle,

could react with the hot 02-H20 propellant exhaust products.

As will be seen later, there was no evidence of combustion.

The observed failure to ignite may be attributed to thi low

exhaust temperaturewhich presumably is inadequate to support

supersonic combtaLion (22). The detailed data are included

in Table II and a summary plot is given i Fig. 4. Ambient

temperature injection through a 0.0625-inch diameter orifice

has been used throughout. For reasons to be presented later,

a modified correlating parameter, (1 + Y )1 /1 s, has beenused in Fig. 4 where, with the exception of the argon data,

quite a good correlation has been provided. go far, all

attempts to locate errors in the argon results have been futile.

- 11 -

The Jah"u N"e~in UAntyAPPOEt PNIunoG LAWRATORY

NNWyg SpOng. M#iyI.W

The reason, if any, for t..is disparity hus not been reconciled.

Characteristics associated with subsonic and sonic injection

are similar to the data presented in Fig. 3.

A series of experiments was undertaken to determine

the gain in axial motor thrust as a result of secondary in-

jection. Ambient temperature injection of CO2 through a

0.180.inch-diameter orifice was used. The axial thrust change,

AF, was measured as a function of WV and F.. The ratio

AF/( FNtan O)was computed and found to be 1.30±0.09, which was

independent ol ; within the accuracy of the experiment.

Since the pressure rise associated with secondary injection is

distributed about the circumference of the nozzle and FN is

the integrated force component in the plane containing the

orifice and nozzle centerlines, a value of AF/(FNtan )larger

than unity (flat plate value) is to be expected.

- 12 -

I*@ Jom'I Hoolir univ..,1yAPPLI90 PHYSICS I"IATORV

INh 11pinip, MAWYiW

V. THEORETICAL FOUNDATION

The foundation for the series of tests reported here

is a modified theoretical model given by Vinson, Amick, andLiepman (8). This model has been found (in general) to be in

qualitative agreement with the bulk of experimental data.

Because of its simplicity and flexibility it served as a guidein selecting experimental parameters. This two-dimensional"weak" jet model assumes that the injected gases expandisentropically and without mixing to form a step-like obstacle

to the supersonic flow. An oblique shock wave which causesflow separation followed by a Prandtl-Meyer expansion is intro-

duced to pr-vide proper flow deflection of the mainstream.

(Fig. 5. )

The net force acting upon the wall as a result of the

injected gases for this model can be simply computed fromlinearized supersonic flow theory (which should be valid ifmainstream deflections are not too large) providing one inte-

grates along the control surface indicated in Fig. 5. Accord-ing to linearized supersonic flow theory (23), the pressure

coefficient Cp - 2(P -P1)/p1yU1 2 is given as

C =2 IAZ (1)CT -1 ) str

-13 -

*AML99 PNOW LAMNATORY

where (dy/dx) str is the slope of the streamline. Integration

along the streamline that divides the injectant and mainstream

gives for the total force normal to the mainstream flow direc-

tion (per unit width)

FN J (P -P) strdX = P1YM2 yo (2)

where y. is the asymptotic displacement of the streamline from

the wall necessary to accommodate the injected gases.

The result given by Eq. (2) has been used by Vinson,

et al (8) to compute the induced reaction which is then

added to the jet reaction to obtain the total interference

force. Such a procedure may be approximately correct for

orifices located near the trailing edge ot the body, but within

the restrictions of the linear theory Eq. (2) alone will give

the total reaction for a body of infinite length.

For a finite-sized orifice, a pseudo two-dimensional

analysis would give

FN = PlA.YMI2/(MI2 -1)1 /2 (3)

where A. is now the asymptotic area through which the injec-

tant passes after being expanded to P Isentropic expansion

- 14 -

The J.Ano Nhe&eu U.".IItyAPPLIl .M hSf LA*SATONv

was assumed by Vinson, et al. The PlA. product can be related

to the injectant flow rate of by

PlA, = W (Toj) /M (J") (4)

where T is the stagnation temperature of the injectant and

m (NJM) is the mass flow function defined as

o Y.-l1/2mi (NJ.) = Nj. i+ M Uj. (5)

where is the injectant molecular weight and R is the

universal gas constant.

Substituting Eq. (4) into Eq. (3) gives for the 4f feu-

tive specific impulse of the injectant

1/2 2Fi () YjN (2 * (6)

s jj= Ojl(M ) (Ui2 .*1)

If we normalize Eq. (6) 1,v dividing by the specific

impulse of the var.'u c haurte, sonic .)et of the injectant,

1 (1 4 Y )(T /M (1) (7)

S., oj J

- 15 -

The Jwh. N"e&M; Unft.qviAPPLIES PVNM L8AT 3RATS

we obtain

3 2(I + Y) -3*

0 (u2 J)12s TRi(j. MI I

(8)

Yl 1 2 1 1+ Y 3 1/(M1 2 -1) 1 / 2 MjV2[2+(Y 2

Although Eq. (8) cannot be expected to apply in detail to the

experiments in question, some general or qualitative interpre-

tations can be deduced that will aid in analyzing or correlat-

ing the data and in making predictions as to secondary in-

jection performauace. Some of these features are:

(1) As long as one assumes an adiabatic process

for the injected gases, the parameter

(1 + Y )Is/Is* is independent of i and T

and essentially independent of V..

(2) Within the restrictions of the linear theory,

the extent of boundary layer separalion does

not affect the magnitude of the interference

force resulting frnm secondary injection.

However, tne pressure rise associated with the

induced shock (which by postulate gives rise to

the separated flow) can be expected to influence

the thermodynamic process of the injected gases.

i.e the extent of total pressure loss (if any)

- 16 -

APP OI PNVGW LANNSIAVme~o, AMeoyI

an a result of possible ihock formation in thejet gases.

(3) The mainstream values for Y and (more uignifi-

cantly) MI appear in Eq, (8) and can be expected

to influence secondary injection performance,

whereas the molecular weight A and temperature

T of the mainstream do not appear to be signifi-

cant parameterm,

(4) The pressure ratio Poj/P 1 will determine Mi.

for any given mainstream condition%; however,

UM cannot be calculated R priori mince the

thermodynamic process of the jet games can be

expected to depend upon the separatod flow

conditions, iince M will become amall am

Pei /PI approaches unity, I /T should I4ncreaso

with decreasing P oj/P and, in fact, will dkverge

at Mj ;- 0.

Because of the several restrictions on this theory, it

cannot be used directly for the analymis of secondary usa in-jection in rocket nozzle flows, 'rhe qualitative arguimunts

presented above nevertheleam can still be expected to he valid

- 17 -

?h JA W~i M IJAUvlrSttyAPPULIS PNVI IAN AIGY

VI. DISCUSSION OF THtE DATA

The data presented in Fig. 3 on the effect of injectant

orifice diameter were taken in order to examine the theoretical

postulate that the effectivenssm of secondary injection should

increase as the PoJ/PI ratio is decreamed. This diverging

characteristic has been observed in a number of experiments

reported by others, but the low prsmiu.e r tio extremes have

not been critically examined. The sonic injection data pre-

sented in Fig. 3 show the characteristic ducline that has also

been observed by others in secondary injection performance for

increasing Po, /P ratios.

However, this trend does not prevail for subsonic in-

Jection; it has been observed that as the .Jet becomes xubhonic

(an indicated by the knee in the disc' rgv coefficient curve),

performance tends to decrease slighti) with docroa ing Po~j/Pl.

The most efficient performance is achieved at or near the

transition from sonic to subsonic injeutLion. This behavior

is not unique to this experiment (15). It in worthwhile to

point out that this transition prossure ratio (P oJ/PI)tr iscompatible with an effective Jet back prpssure that will give

combined "Just-choked" jot fLow _and * turbulent boundary layerseparation of the mainstream ahead of the port. The pressure

rise to give turbulent boundary layer iiwparation for M* , 2.4

air is approximately (P2/P )Hup , 2.2 (24). If one assumN

that the static pressure at the Jet orifice P is approxi-

mately equa. to the presure in the separated region P2, tho,

(~y",Pltr r. (Po/P)(P2/P,,, ,, (2.0)(2.2) 4.4

which is iii rvaoniable agreement with Fig. 3.

- 18 -

tih e $ qAIl .UvIe4idAPPllLIN PNIOUI iLllNlliON

Additional information on the pressure rise associated

with the induced shock wave can be obtained from the subtionic

injection data in the following manner: If one assumes one-

dimensional isentropic flow for the jet games and P discharge

coefficient equal to its asymptotic value at large PoJ/Pit

the static pressure at the jet orifice P can be evaluated

from the measured values of Wj, P UP ToJ' and Aj, In addition,

we assume that P - P2 . The results of such calculations for

the subsonic CO2 injection data appear in Fig, 6. For Poji PI

less than about 3, the pressure rise P2 /P1 increases almost

linearly with P oj/P (increasing jet flux), For Po /Pl greater

than about 3, P2 /P1 tends to level off at approximately a

value required to give turbulent boundary layer separation.

This would suggest that the induced oblique shock wave is

probably attached initially to the leading edge of' the orifico,

and its strength increases with increase in jot flux until

it roaches a limiting pressure rise sufficient to give sopara-

tion. The shock then detaches from the orifice lip and moves

upstream with increasing PoJ/p l . This interpretation is comn-

patible with the theoretical model discussed abovu.

The theoroLlcul mcdel fails, however, to present oveni

a qualitatively correct interpretation ol the subsonic injuc-

tion data for which it should be most applicable: i.e.. it

dos not predict a decreasing I /Is for decreasing P /P.

Thib characteristic has not booi demonstratud for any kiown

theoretical descriptlon.

In addition, the strong influelnce that d has upon the

effectiveness of secondary gas injection as showi, iy the data

in Fig. 3 had not been established in the reports ol other

- 19 -

1he Johm HOW..i VANvWeIltyAPPLIID PNVSIW LANRATIAV5I~w Siql~, W4.?Iud

experiments and was not anticipated. Theory has not been

developed to the point of including three-dimensional effects

and is of no help in Interpreting these data.

It is quite interesting, however, to replot the data

of Fig. 3 as Is /1 versus * /., which is essentially the

form frequently used by others to report secondary injection

data, Figure 7 is much a graph and shows that all sonic CO2injection data correlates rather well, (The subsonic injection

data do not correlate on this plot.) The reason for the corre,,

lation is not clear since the independent variable W j/

suggests a kind of one-dimensional flow not physically plausibte

or consistent with restriction of the induced oblique-shock

pressure rise (approximately) to within the Mach cone emanat-

ing from the wall-jet perturbation. Nevertheless, a one-

dimensional model with assumed complete mixing of the injectant

gases with the supersonic flow has been given by Bonham andGreen (25) to establish the releiant parameters for secondary

injection. Such an approach fails, however, to predict. theeffect of variations in injectant molecular weight observed

experimentally in this study. It muitst be concluded that the

effect ot d on secondary gas injection is not welt under-

stood and a more comprehensive analytical model would be

welcomed.

With the correlation provided in Fig. 5, rolative

changes in injectant molecular weight and specific heat ratio

appear predictable with Eq. (8). Because the H 2 injection

data appear to correlate well witn other inert gas injection

data, it has been concluded that no combustion takes place

between the H2 And the hot H20-02 exhaust gases. This c.on-

clusion in consistent with the research of Chinitz and Gross

(22), who reported that combustion between H2 and heated

-20°

the WA.m HOP4106 URIV916ItvAPPLUS PM14110 #AMATSV

supersonic air does not occur below a critical stagnation

air temperature of about 20000P. By inference, the relative

effect of changing injectant total temperature can also be

predicted, but additional data on this parameter would be

desirable. These conclusions unfortunately must be qualified

somewhat because of the nonconforming but apparently error-

free argon data.

Additional experiments designed to measure the effects

of other parameters (such as, Toj, T0, M1 , Pl, and nozzle

geometry) are desirable in order to establish appropriate

theoretical evenues. Some of these experiments are presently

in progress at this Laboratory.

- 21 -

7

Me JuIsoo "OolhA Volvo's-IVAPPLIS PNV81lS LANNATIOYSilve so", MetYlisd

Table I

Properties of research rocket motor, nosle, and Injectantsused in secondary gas injection experiments

Motor

Propellant O% U20 23xhauet gs romposition (0.70. mole fraction H20

(0.292 mole fraction 02Thrust coefficient, C? 1.42

Prsure, P. - 400 lbs/in 2

Propellant flow rWil, - 0.84 lbm/eoc

txbq.qt gas total temperature, TO 1845R

Rpeit:fic heat ratio, Y 1.266

Ambient pressure atmospheric

InljectantGas CO2 0 N2PHe, He + Ar, Ar, and H2Injectant pressure, Poi 40-1000 lbs/in2

Injectant total temperature, Toj -70'r

Injectntnt port diameter, d 0,0625, 0,089, 0.125, 0.180 in.

Nouzle (Conical, nharp-edged throat)

Divorment half Anglo, n 15 degrees

Throat diameter, dt 0.501 Inches

XxIt diameter, d,, 1,074 Inches

Nozzle diamotr at. injectant port,dI 0.812 inches

Mach Number at injectlon planv, M1 2.34

xit Mach Number, M3 2.03

- 22 -

The JeA M4o.&,., Urn,4.,Itv~APPt, I PagOg bOUW Aoaov

UK*w Sol". MeeVIe4

Table I1

lecondary Oa Injection Data

Injectant - CO2 Orifice Diameter, d- 0.180 inches

PO Poj /P 1 X / /1s b

t: D C

paia pala iba/mec, ib., *Cc.

393 48.3 0.0088 0.52 86.2 1.57 1.90 0.186

398 52,3 0.0086 0.80 90.7 1.78 2.00 0.238

402 83.3 0.0088 0.84 94.0 1.81 2.08 0.327

410 89.0 0.0120 1.12 92.5 1.97 1.04 0.288

409 63.3 0.0143 1.39 93.2 2.R 2.06 0.331

404 66.7 0.0175 1.54 87.8 2.26 1,94 0.367

406 72.0 0,0207 1.95 92.9 2.43 2,05 0,399

400 75.4 0.0230 2.13 92.6 2.50 2.04 0.426

4)7 81.1 0,0292 2.61 80.4 2.73 1.97 0.802

4o16 80.7 0.r288 2.64 89.9 2.74 1.99 0.807

310? 91.6 0.0423 3.75 88.2 3.15 1.95 0.648

408 102 0,0584 4.86 88.4 3.43 1.95 0.747

412 110 0.0640 3.83 87.7 3.66 1.94 0.799

393 108 0.0600 .64 87.1 3.71 1.92 0.835

402 110 (), (M N 3.57 .4.1.1 3.73 1.84 0.850

403 131 0.0898 7,319 84.5 4.45 1.87 0,913113 138 0.0907 7.99 84.8 4.5f8 1.87 0.944

J!18 141 0.0991 8.111 84.1 4.86 1.86 0.945

- ,xI -

SAoh. t4WIf Unof oqgI ,APPISSI PUVUU t4WBIASPII

Table It - continued

.zJectant -CO 2 Orifice Dimeter, dj - 0.125 inches

o o j 1F1* a P j/P1 I.a .b CDc

pain pNta lbs/uec, lb., mec.

398 67.0 0.0066 0.82 92.6 2.28 2.04 0.361

397 76.0 0.0130 1.25 94.0 2.58 2.08 0.503

403 86.2 0.018 1.83 96.3 2.92 2.13 0.638404 87.6 0.0213 2.03 94.4 2.98 2.08 0.711

404 96.2 0.0256 2.47 95.7 3.28 2.11 0.778

412 105 0.0208 2.93 98.3 3.48 2.17 0.824404 105 0.0307 3.02 97.2 3.55 2.15 0.853

407 120 0.0379 3.79 100.6 4.01 2.22 0.902

404 142 a1.0462 4,47 97.6 4.82 2.15 0.925408 166 0.0547 5.32 97.9 5.59 2.16 0.938

400 165 0.0553 5.17 94.7 .1.60 2.09 0.942402 187 0.0622 5.85 93.8 6.36 2.07 0.951

403 202 0.0693 6.32 92.8 6.86 2.05 0.953

- 24 -

A9PKIIIN t"N LANAIOVIm -n. -q oq

Table II - continued

Injectant - CO2 Orifico Diameter, dj - 0.089 Inches

P O P o j i j 'i, I1 1/ p 1/ 1 * b C D

psla psia lbm/sec. lbs. sec.

406 196 0.0316 3.27 103.6 6.57 2.29 0.911

402 201 0.0331 3.30 100.5 6.77 2.22 0.912

401 264 0.0435 4.29 98.6 8.92 2.18 0.915

405 273 0.0439 4.33 98.5 9.2 2.18 0.892

420 326 0.0541 5.25 98.1 11.0 2.17 0.901

397 328 0.0539 5.10 94.6 11.2 2.09 0.905

402 356 0.0581 5.58 95.7 12.1 2.11 0.898

401 358 0.0600 5.70 94.9 12.2 2.10 0.916

- 25 -

The )OKAS HkIb. Unl.erl s

APPLICR PeNVICS LAISRATOR-

Table I - continued

Injectant - CO2 Orifice Diameter, d - 0.0625 inches

Po oJ a PJ/P 1 1/ISb CDC

psia paln lb/mec. lbs. sec.

401 91.3 0.0057 0.64 107.5 2.98 2.37 0.756

401 102 0.0077 0.89 112.7 3.33 2.48 0.899

400 128 0.0098 1.11 110.9 4.20 2.45 0.903

399 159 0.0116 1.35 115.5 5.23 2.35 0.845

401 178 0.0144 1.69 116.7 5.94 2.5.1 0.929

393 186 0.0132 1.60 118.0 6.23 2.61 0.831

395 211 (1.0152 1.79 115.f 7.10 2.55 0.839

399 246 0.0186 2.13 114.3 R.13 2.52 0.864

399 275 0.0223 2.51 111.9 9.24 2.47 0.923

389 309 0.0328 2.62 109.1 10.6 2.41 0.878

396 354 0.0268 2.94 109.1 11.8 2.41 0.851

411 387 0.032q 3.63 111.3 12.7 2.46 0.933

400 388 0.0298 3.22 106.7 12.8 2.36 0.862

395 413 0.0316 3.23 102.4 13.8 2.26 0.843

400 464 0.0356 3.69 102.7 15.4 2.27 0.836

- 28 -

the eha KM.ns U.Rf.''syAPPLI9 PNVIIOS LASBAIORY

SINW IPun. #A'WyIe

Table I - continued

Injectant - CO2 Orifice Diameter, d- n 0.040 inches d

Po Po jF N 1 P/Pl Iu b C

P. po Iaa poil 15O CDcpn;In pnIa ibs/sec. tbs. sec.

395 215 0.0056 0.68 1!8.9 7.19 2.63

399 277 0.0073 0.91 121.6 9.26 2.68

397 ?61 0.009e 1.16 118.8 11.6 2.62

398 398 0.0110 1.27 114.2 13.3 2.52

4 05 *1)3 0.0141 1.65 114.5 16.7 2.5-

- 27 -

the JIw4 N&IMm U01vweidyAPftIKe PHNtesv LABORATORV

Table 11 - continued

Injectant - Ar Orifice Disveter, dj - 0.0625 Inches

po 3 ;N 1 " a/PI 1/1,b cV

pala Pala lbs/ec. b. sec.

393 65.5 0.0028 0.37 127.1 2.22 2.84 0.498

392 74.7 0.0041 0.84 128.4 2.54 2.87 0.617

393 75.3 0.0042 0.52 119.4 2.55 2.67 0.646

389 84.8 0.0056 0.66 119.2 2.90 2.67 0.760

405 90.7 0.0059 0.76 122.8 3.03 2.75 0.769

391 101 0.0073 0.4 114.5 3.42 2.56 0.817

3S 120 0.0091 1.12 120.0 4.02 2.69 0.870

391 123 0.0091 1.14 123.9 4.21 2.77 0.836

411 140 5.0101 1.32 127.1 4.63 2.84 0..36

39 158 0.0119 1.48 121.6 5.34 2.72 0.862

400 167 0.0129 1.55 120.5 5.62 2.70 0.662

399 164 0.0144 1.82 125.9 6.13 2.82 0.882

402 186 0.0134 1.72 122.41 6.24 2.73 0.854

405 221 0.0162 2.01 124.S 7.46 2.79 0.821

407 220 0.0171 2.08 119.9 7.64 2.68 0.855

404 226 ".0171 2.09 118.1 7.68 2.64 0.870

399 246 0.0139 2.26 121.0 8.72 2.71 0.862

399 285 0.0220 2.59 114.2 "56 2.56 0.89402 317 0.0255 3.04 116.8 1"..5 2.66 0.907

404 34' 0.0271 3.14 113.7 11.5 2.54 0.896

396 394 0.0303 3.42 112.1 13.3 2.51 0.870

400 409 0.0323 3.60 111.1 13.7 2.48 0.889

404 552 0.0434 4.21 97.6 18.4 2.18 0.880

406 660 0.0544 5.31 98.0 22.2 2.19 0.896

401 895 0.0721 6.51 91.5 29.9 2.05 0.885

-28 -

the j.Ihes moofl4 Uff...'sI,APPkIIID PSONICS LAOSPASIV

Table II - continued

Injectant - M2 Orifice Diamter, - 0.0625 inches

P. poi a F I a a/II*b C

psla ps. lbrm/sec. lbs. sec.

393 110 0.0061 0.82 133.5 3.69 ".43 0.804

139 0.0080 1.06 129.3 4.68 2.36 0.858

397 171 0.0102 1.36 129.5 5.73 2.36 0.63

392 213 0.0127 1.70 130.1 7.16 2.37 0.879

397 266 0.0166 I..16 128.4 8.84 ..34 0.904

397 325 0.0200 2.46 121.9 10.9 2.22 0.890

391 440 0.0274 3.39 121.8 14.4 2.22 0.902

391 579 0.0349 4.19 117.4 19.5 2.14 0.883

398 709 0.0435 4.95 114.0 23.5 2.08 0.877

- 933 0.0579 0.26 108.6 31.4 1.98 0.886

- 29 *

APFLSCD PNVSICG LAONIATORY

Table II - continued

Injectant - Ne Orifice Diameter, d- 0.0625 inches

p6 p 0i 1 N Ia a 1j PI 1 /*b CDcPo joi a ojl i a l.* D

Pula Pea lbu/sec. lbs. sec.

389 86.4 0.00173 0.53 297.4 2.03 2.13 0.748

395 ".2 0.00171 0.51 257.3 2.95 2.04 0.731

392 104 0.00224 0.6 253.5 3.51 2.01 0.812

462 120 0.00256 0.78 290.8 3.61 2.06 0.807

399 140 0.00318 1.01 307.7 4.79 2.18 0.851

389 159 0.00376 1.15 297.5 5.41 2.11 0.862

390 201 0.00472 1.41 :989.5 6.85 2.05 0.879

393 242 0.00593 1.77 292.3 8.16 2.07 0.910

390 301 0.00726 2.18 293.1 10.2 2.06 0.896

393 365 0.0009 2.57 283.6 12.3 2.01 0.902

394 459 0.0110 3.03 272.8 15.0 1.93 0.879

409 656 0.0166 4.40 267.5 21.4 1.89 0.930

399 933 0.0224 5.97 265.6 31.9 1.88 0.875

- Jo -

the OAS NoD.01 UA wIIllyAPPLaII PeiYPIOG %A004*?SIY

1,1w, Ipfl~, MAJyand

Table It - continued

Injectant - He + Ar Orifice Diafeter, dj - 0.0625 inches

PC po0 i r I I J /P I/I. b C c

pmia ps1a lbm/eec. lbs. sec.

393 90 0.0033 0.63 185.2 3.03 2.20 0.1800

395 113 0,0044 0.86 190.2 3.90 2.25 0,829403 130 0,005) 1,07 192.2 4.53 2.23 0.866402 178 0,0072 1.35 137.0 5,96 2.22 0,859

404 225 0.000 !.62 184.3 7.42 2.18 1.931403 307 0.0136 2.50 183.3 10.2 2.17 0.943

403 432 0.0181 3.00 170.3 14.3 2.02 0.89..02 853 0,0242 3.95 164.1 13.4 1.94 0.92P401 709 0.0303 4.83 160.4 21.5 1.90 0.9400 906 0.0394 6,10 157,9 30.2 1,67 0,919

3- -J

APPLISI PNAIGI, %A140A9OV

Table I - continued

Znjotast - H Orifice Diameter, d3 - 0.0605 inches

ip0 pel i CD0

0 Poi j FN t aA Po/P 1/1 eb

pas pit& lbe/soc. lbs. e.

403 101 0.00170 0.33 478,0 3.68 2,34 0.86

394 123 0.00336 1,00 470.1 4.17 3.33 0,96395 170 0.0016 1,4? 461,4 5,5 3.30 0.965

413 235 0.0043 1,l 413.4 7,00 3.11 0.647'01 343 0,00013 3.4 411.5 11.5 3.01 0.9664(11 444 0.0003 3.45 416 14.6 3.05 0.675326 599 0.0119 4.71 390.5 19.7 1.65 1.030401 I0? 0,01517 0.22 M6 36.1 1.65 0.64?

a. Is reduced to 700F by: Is * U (IP )(83@/? W9

b. lte 1q. (7) for definition of Is

* (T )l/2 3D c, A j o l where it 1 -0.00770 (P/T ) +.09 (oj/To) i

a compremnibtlity correction (18) used for C0 data - 1, for the otherlaie.t),

d. 0.0025 inch diameter orifice partially plugged with solder; exact diameterunknown.

- 32 .

&IIunm PmVII" LAwmavep

Table III

lame Plaoertiee of the Xne@tants Used inSecondary Gas Inject ion Studies

GA1 Yj is (01C)

CO 2 44.01 1.30 45.3

N228.02 1.40 54.9Ar 30.91 1.67 44.70.3 He + 0,2 Ar, 21.18 1.67 84.4me 4.00 1.67 141.2H2 2.02 1.40 204.8

a Mole fractions

Yh. Jie~eNq~ge UAVwit

idi

APPM I I PNIIII &ABINA10MV

IMI

imi

I ~ 5I

[I ,l/1 3

El+i 0 Li

i'3 6- :31 -

APpugs PNWue OAgNAysm

lik low #4 i

4.r*48b04 0

a 0 0

0 Boo

0

V .C

'In

* ~ 4 -

II

Tiw Ji.~,Mwh,,Ue~irodt

Ap~bgs PyOWS~ASS4T@L

*ppI, ie eNyeica IAUITOSRY.tvw erling, M* yI~d

/I dj |nks

• 0.120

!6

_. 2

//-- SONIC JET

Poj/Pl

Fig. 6 APPROXIMATE PRESSURE RISE DUE TO INDUCED SHOCK WAVE(SUBSONIC CO2 lMMJRCTWN DATA)

- 39 -

3.0 -1

+ y+ ft 0.04(0.0615 Inch@efepo,*gegy

/ *Igomet Unknown

A(

W4fo A2- d ft af

0 %, 44.

d~b __J

fig. 7 CORRELATION OF SONIC SECONDARY INJECTION DATA(CO2 INJXCTANT AT 70Vf; ALL SUBBOXIC

* INJECTION DATA (Poi/P1 4. 0)ARE PLAGGID

-4o

fte JoA Hmooi UnivorityAPPLIII P11116111 L*SfAOYt

REFERENCES

1. United Aircraft Corporation Report R-63143-24, ThrustAxin Control of Oupersonic Nozzles by Airjet Shock- Inter-erence, y 0. F. iHausmann, 2 MaY l''2.

2. WADD Technical Report 80-329, Interaction Effects ofSide Jets Issuing From Flat Plates and Cylinders AITnedMTE a K0_uroj3 n at r'#ai -76Y7 J- L C~andP.B. Hyway -l10.

3. Liepman H. P., "On the Use of Side-Jets as Control Devices,"ARS Journal, Vol. 29, June 1959, pages 453-455.

4. NASA Technical Note D-649, Loads Induced on a Flat-PlateJet bTroughWief by an Air Jet Exhausti Tni-Flrendtcuarg -u

- ng-and-Normal to a Free-Stream Flow of MachNu mbe r-. b7y7 7 nos,-V arch 1961T -

5. NASA Technical Note D-580, Surface Pressure DistributionsWith a Sonic Jet Normal to Adjacent Flat Surfaces at Mach2U o -o- -7 R TW uSiio-sn-B7 H. Andersn-,and'J.WaIrd, "eb-ruary 1961.

6. NASA Technical Note D-743, namic Interaction EffectsAhead of a Sonic jet Exhaustng r endicularly From ala t Pl-t a Mach Number 6 Free Stream, by D. J.

Romeo and J_.T i erett, April i96.

7. University of Michigan, Department of AeronauticalEngineering, WTM-255, An Experimental Investigation ofthe Forces and Flow FieTd Produced by a Jet Exhausting-teT FFrom a ne- li nder in a-a-ch-28 8 -team, by

J~t7 ic~~7ETndandH7P.LiemaiTo~mez'1955

8. NASA Memo 12-5-58W, Interaction Effects Produced by JctExhaustinyg Laterally Near Base o TUgive-Cy1inder Modelin- SersonicMainlteam. b y-P.--7W .Vins i-J.L. Am-ic..k,andd If.P.Tipman Fe-iiiry 1959.

9. University of Michigan CM-979, Jet Interference ExpLrientsEmploying Body-Alone and Body-F-inin Configurations at

~ c~ds TVb-. .--- arTvalho and P. B. Hays,December 1960.

-41 -

the Johw Hwhk',.no UerglyAPPI|N PSyNll kAbIASIDRY

4IF4w Itng, wI ld

10. United Aircraft Corporation Roport R-0937-33, Jet-InducedThrust-Vector Control Applied to Nozzleo Havin-_TaW --My R. ... .

11. NAVORD Report 5904, Vol. 13, A Theoretical. and Exveri.-mental Investigation of a fethod of Thrust Vector coNtrolfor Old ocket Motors, by U',-lhahr- -n "Tdwards,

Decemer W W

12. NASA Technical Memorandum X-416, A Cold Flow Investiga-tion of Jet-Induced Thrust-Vector-Cot-ro ,' y J. E.-WcTiy anT . aTF= eZ .,.6iF TF.-

13. Bulletin of the 17th Weetingt, JANAF-ARPA-NASA Solid Pro-pellant Group, Vol. I1, An Experimental Investigation ofJet-Induced Thrust Ve ' I.r"Control Methods, by C. J.

14. Bulletin of the 17th Meeting, JANA'-ARPA-NASA Solid Pro-pellant Group, Vol. III, Propel'ant Gas Injection forThrust Vector Control of 8ol'Vroeilant -u-kewk byNbT-TDrewry, IT-RinZ-tiuk T. E. XallmeyerF_71ITHarmoning, D. P. Hanley and D. P. Hug, May 1961.

15. Paper presented at Fifth Symposium on Aeroballistics atU. S. Naval Ordnance Laboratory, Research on SecondaryInjection for Thrust Vector Contri plations, byR.O.Sliate s, K7JGer-, C-T nham, and F.Johnson, 16-18 October 1961.

16. ARS preprint No 2216-!1 of paper presented at ARS SpaceFlight Report to the Nation. New York Coliseum, An Experi-mental Investigation of Shock Vector Control With GaseousSeconfv t inn, T)-yE7 -ad C -y Adi-A-do-rquist,

17. JHU/APL Bumblebee Report No. 286, JIterforenco Between aJet Issuing Laterally From a Body aad the Xnveloln1g'Sup'ersonic Stream, by L.r.Fr--, -;''-9"-

18. Jain-Ming Wu, R. L. Chapkis, and A. Mager, "Approximat,Analysis of Thrust Vector Control by Fluid Injvcttion,"ARS .Journal, Vol. 31, pages 1677-1685.

19. ARS preprint No. 2335-62 of paper presented at ARS SolidPropellant Rocket Conferen(ce, Baylor University, AnAnalysis of Gaseous Secondary Injection Into Rocket"

- 42 -

The JoIwo Homkins Ui ivetilyAPPLIgs PHY9e IANINATRY

111n, la toI

20. Becco Research and Development Department Bulletin No. 67,Hydrogen Peroxide Physical Properties Data Book, 1955.

21. NACA Report 959, One-Dimensional Flows of an ImperfectDiatomic Gas, by .T-J. Eggers, Jr.,-T IT5r.

22. Fairchild Engine and Airplane Corporation for ProjectSQUID (ONR), Contract No. NR 1858(25) NR-098-038,Exploratory Studies of Combustion in Supersonic Flow, byW.-Chinitz and R. A. -rossi June '11-59.

23. Shapiro, A. H., T.e Dynamics and Thermodynam-cs of Com-pressible Fluid Flow, The aonaid Press nC., N 0rk,-T953,

24. NACA Report 1356, Investigation of Beparated Flow inSupersonic and Subsonic Streams wit asis on Transi-tion. :y D.-R. Chapman, D. T. Ruehn, and H. K. Larson,

25. Bulletin of the 17th Meeting, JANAF-ARPA-NASA SolidPropellant Group, Vol. 11, Parameters Controlling thePerformance of Secondary Injection, by C. B.Benham andC. J. Green,-ayT1961.

- 43 -

The J.Ihu I4&MM UMWn~tyAPittt Fp M MIAryRd

ACKNOWLEDGEMENTS

The authors would like to express their apprecia-

tion to E. Schmidt and J. Loveless of the APL/JHU High

Temperature Laboratory for their assistance in the

Laboratory, to I. Soslow for preparation of drawings

and figures, and to R. H. Cramer for his continuing

interest during the course of this investigation.

- 44 -

The JAP, '00kino UA.Veif~yAPP1OIS P, lO0 LAWRATIIV

SIee &VpFg, #ueYle.W

The technical papers and progress reports issued

by APL in the CM series are characterized by extensivo

treatment of their subjects. Official Laboratory review

of CM reports substantiates their technical validity and

establishes suitability for distribution to qualified

personnel outside Section T.

In addition to internal (Section T) distribution,

initial distribution of CM-lOO has been made in accord-

ance with Guided Missile Technical Information Distribu-

tion List MML 200/23, List No. 23, dated 3 April 1961.

- 45 -